GOE 6K AIRFOIL (goe06k-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 6K AIRFOIL (goe06k-il) Reynolds number: 100,000 Max Cl/Cd: 45.65 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe06k-il-100000.txt Download as CSV file: xf-goe06k-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 6K AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4860 0.11086 0.10608 -0.0337 1.0000 0.0588
-8.500 -0.5003 0.10919 0.10449 -0.0333 1.0000 0.0589
-8.250 -0.5136 0.10721 0.10258 -0.0333 1.0000 0.0589
-8.000 -0.4782 0.09884 0.09415 -0.0276 1.0000 0.0632
-7.750 -0.4857 0.09658 0.09195 -0.0260 1.0000 0.0644
-7.500 -0.4972 0.09454 0.08998 -0.0242 1.0000 0.0664
-7.250 -0.5078 0.09221 0.08771 -0.0232 1.0000 0.0671
-7.000 -0.5168 0.08948 0.08503 -0.0241 1.0000 0.0699
-6.750 -0.5218 0.08649 0.08205 -0.0254 1.0000 0.0709
-6.500 -0.5263 0.08387 0.07929 -0.0301 1.0000 0.0723
-6.250 -0.5269 0.07890 0.07441 -0.0281 1.0000 0.0742
-6.000 -0.5237 0.07587 0.07144 -0.0248 1.0000 0.0768
-5.750 -0.5199 0.07277 0.06830 -0.0244 1.0000 0.0801
-5.000 -0.4992 0.06211 0.05737 -0.0255 1.0000 0.0923
-4.750 -0.4853 0.05880 0.05362 -0.0277 1.0000 0.1014
-4.500 -0.4768 0.05543 0.05035 -0.0256 1.0000 0.1053
-4.250 -0.4629 0.05227 0.04693 -0.0259 1.0000 0.1169
-4.000 -0.4491 0.04957 0.04403 -0.0254 1.0000 0.1303
-3.750 -0.4354 0.04695 0.04130 -0.0245 1.0000 0.1446
-3.500 -0.4220 0.04450 0.03884 -0.0231 1.0000 0.1615
-3.000 -0.3994 0.04073 0.03510 -0.0193 1.0000 0.2348
-2.250 -0.2974 0.02963 0.02114 -0.0176 1.0000 0.0836
-2.000 -0.2732 0.02739 0.01825 -0.0157 1.0000 0.0749
-1.750 -0.2520 0.02604 0.01674 -0.0145 1.0000 0.0789
-1.500 -0.2291 0.02479 0.01519 -0.0131 1.0000 0.0807
-1.250 -0.2059 0.02373 0.01388 -0.0118 1.0000 0.0820
-1.000 -0.1831 0.02293 0.01286 -0.0105 1.0000 0.0872
-0.750 -0.1615 0.02210 0.01209 -0.0095 1.0000 0.0975
-0.500 -0.1404 0.02133 0.01135 -0.0081 1.0000 0.1064
-0.250 -0.0427 0.01835 0.01129 -0.0223 1.0000 1.0000
0.000 -0.0219 0.01866 0.01118 -0.0211 1.0000 1.0000
0.250 -0.0016 0.01899 0.01126 -0.0202 1.0000 1.0000
0.500 0.0187 0.01934 0.01140 -0.0193 1.0000 1.0000
0.750 0.0390 0.01971 0.01161 -0.0184 1.0000 1.0000
1.000 0.0591 0.02009 0.01187 -0.0176 1.0000 1.0000
1.250 0.0792 0.02050 0.01218 -0.0169 1.0000 1.0000
1.500 0.0992 0.02093 0.01253 -0.0161 1.0000 1.0000
1.750 0.1190 0.02139 0.01291 -0.0154 1.0000 1.0000
2.000 0.1388 0.02187 0.01334 -0.0147 1.0000 1.0000
2.250 0.1841 0.02285 0.01430 -0.0194 0.9904 1.0000
2.500 0.2261 0.02379 0.01525 -0.0232 0.9818 1.0000
2.750 0.2661 0.02459 0.01608 -0.0266 0.9730 1.0000
3.000 0.3005 0.02512 0.01666 -0.0288 0.9629 1.0000
3.250 0.3360 0.02569 0.01732 -0.0312 0.9528 1.0000
3.500 0.3735 0.02628 0.01804 -0.0338 0.9428 1.0000
3.750 0.4155 0.02687 0.01877 -0.0371 0.9328 1.0000
4.000 0.4569 0.02726 0.01933 -0.0402 0.9209 1.0000
4.250 0.5065 0.02692 0.01922 -0.0438 0.8984 1.0000
4.500 0.6397 0.02261 0.01557 -0.0576 0.8515 1.0000
4.750 0.7291 0.01597 0.00950 -0.0591 0.7405 1.0000
5.000 0.7536 0.02012 0.00967 -0.0565 0.1077 1.0000
5.250 0.7661 0.02149 0.01094 -0.0535 0.0870 1.0000
5.500 0.7886 0.02333 0.01273 -0.0522 0.0767 1.0000
5.750 0.8258 0.02572 0.01505 -0.0539 0.0664 1.0000
6.000 0.8644 0.02805 0.01755 -0.0556 0.0604 1.0000
6.250 0.9049 0.03131 0.02103 -0.0574 0.0585 1.0000
6.500 0.9371 0.03600 0.02602 -0.0581 0.0568 1.0000
6.750 0.9519 0.03725 0.02782 -0.0547 0.0550 1.0000
7.000 0.9762 0.04196 0.03280 -0.0538 0.0572 1.0000
7.250 1.0010 0.04758 0.03990 -0.0476 0.0986 1.0000
9.250 0.9732 0.07981 0.07468 -0.0171 0.1350 1.0000
9.500 0.9083 0.07876 0.07390 -0.0101 0.1365 1.0000
9.750 0.8732 0.08059 0.07587 -0.0049 0.1370 1.0000
10.000 0.8393 0.08400 0.07936 -0.0020 0.1371 1.0000
10.250 0.8065 0.08850 0.08392 -0.0014 0.1372 1.0000
10.500 0.7714 0.09414 0.08959 -0.0031 0.1374 1.0000
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